Institute of Aerodynamics


Institute of Aerodynamics

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Geissler W.,Institute of Aerodynamics | Wall B.G.V.D.,Institute of Flight Systems
Annual Forum Proceedings - AHS International | Year: 2010

As early as 1934, Prof. Dr. Hans-Georg Küssner from the Aerodynamische Versuchsanstalt Göttingen (AVA) in Germany was thinking about an alternative to helicopter rotor torque compensation, that was tried experimentally already in 1921 by Passat in England. Based on the theoretical work about two-dimensional unsteady airfoil aerodynamics, especially combined plunge and pitch of an airfoil that revealed a propulsive force as is used by bird flight, he was tempted to try this concept on driving a helicopter rotor via forced flapping of its rotor blades. Experiments have been carried out in the mid-30ies and patents applied for the concept which were granted a long time later after an intense and controversy dispute with the responsible officials. This paper describes Küssner's efforts regarding the forced flapping concept and experimental results he obtained in the AVA wind tunnel in hovering condition. © 2010 by the American Helicopter Society International, Inc.

Thouault N.,TU Munich | Thouault N.,Institute of Aerodynamics | Breitsamter C.,TU Munich | Breitsamter C.,Institute of Aerodynamics | And 7 more authors.
AIAA Journal | Year: 2012

The aerodynamic characteristics of a rotating cylinder in crossflow are investigated by means of unsteady Reynolds-averaged Navier-Stokes simulations. For a cylinder configuration with endplates, the numerical simulations match the experimental trend for the force coefficients and the Strouhal number. Design parameters are studied including the ratio of circumferential cylinder velocity to freestream velocity (α), the endplate diameter ratio, and the cylinder aspect ratio. The incoming flow separates on each endplate edge and rolls up into two tip vortices that merge downstream. They impact considerably on the configuration performance, particularly at high α. The tip vortices influence the cylinder flow topology, especially for low-aspect-ratio cylinders with small endplates. Finally, a cylinder configuration with spanwise disks is investigated forα<3:4. The streamwise velocity component increases between the boundary layers of two facing disks, thereby decreasing the effective spinning ratio. At the corner, the cylinder boundary-layer thickness is reduced due to the radial flow component occurring on the disk. Furthermore, adding spanwise disks decreases the strength of the tip vortices. The combination of these three effects leads to a drag reduction at high α compared with a cylinder configuration without spanwise disks. © 2011 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

Hemchandra S.,Georgia Institute of Technology | Hemchandra S.,Institute of Aerodynamics | Peters N.,RWTH Aachen | Lieuwen T.,Georgia Institute of Technology
Proceedings of the Combustion Institute | Year: 2011

Predicting the ensemble averaged heat release response of a turbulent premixed flame to acoustic forcing is a fundamental problem associated with understanding combustion instabilities. This paper describes an analysis of this problem, by modeling the response of a flame that is simultaneously perturbed by broadband, turbulent fluctuations and narrowband, acoustic fluctuations of amplitude a and T, respectively. It is shown that the response of the flame surface to coherent forcing and turbulent fluctuations are coupled even at linear order in coherent forcing amplitude, a, due to flame propagation (kinematic restoration). This coupling effectively causes the local consumption and displacement speeds of the flame to vary in time over a forcing cycle. Turbulent fluctuations also provide a mechanism for destruction of coherent flame surface wrinkling at first order in a, an effect which only occurs at O(a2) for laminar flames. © 2010 Published by Elsevier Inc. on behalf of The Combustion Institute. All rights reserved.

Steimle P.C.,Eurockot Launch Services GmbH | Karhoff D.-C.,RWTH Aachen | Karhoff D.-C.,Institute of Aerodynamics | Schroder W.,RWTH Aachen | Schroder W.,Institute of Aerodynamics
AIAA Journal | Year: 2012

The unsteady transonic flow around a flexible transport-type swept-wing configuration with the supercritical airfoil BAC 3-11/RES/30/21 is studied regarding the effects of dynamic shock/boundary-layer interaction involving separated-flow areas on the local fluid-structure coupling at unit Reynolds numbers of around 1:5 × 10 7 m -1.Afastresponding pressure-sensitive-paint system using pyrene on anodic aluminum binder is applied to measure the unsteady pressure distribution on the entire wing surface, with high spatial and temporal resolution providing a unique view on the spanwise flow development and self-induced shock oscillations. The aeroelastic response of the wing structure is synchronously recorded by videogrammetric deformation measurement. Flow cases with incipient separation and full-scale separation at Mach numbers of 0.86 and 0.92 are compared.A weak shock/boundary-layer interaction with incipient separation has minor effects on the wing structure, despite the occurrence of large pressure fluctuations, whereas the strong interaction involving shock-induced separation results not only in significantly weaker fluctuations in the pressure field, but also in a strong fluid-structure coupling. Copyright © 2011 by the authors.

Konopka M.,Institute of Aerodynamics | Konopka M.,RWTH Aachen | Meinke M.,Institute of Aerodynamics | Meinke M.,RWTH Aachen | And 2 more authors.
42nd AIAA Fluid Dynamics Conference and Exhibit 2012 | Year: 2012

The interaction of a shock wave with an expanding supersonic turbulent boundary layer at a freestream Mach number Ma = 1:76 in a supersonic combustion ramjet inlet is analyzed using large-eddy simulations. In this context, the phenomena of relaminarization and expansion followed by compression are considered. The results are compared to a Ma = 3 computation at the same Reynolds number and an expansion angle of α = 12°. Considering the relaminarization issue a skin-friction coefficient reduction downstream of the expansion corner of 25% is obtained at Ma = 1:76 and a reduction of 50% at Ma = 3:0. Significant Reynolds stress component reductions occur at both cases and the near wall anisotropy tends towards the one-component limit. Semi-log velocity profiles reveal the formation of a large laminar-like sublayer downstream of the expansion which at the Ma = 3 case is four times longer than at the Ma = 1:76 case. The analysis of the shock wave interaction downstream of the expansion corner at Ma = 1:76 shows an 11-fold increase in the wall-normal Reynolds stress component and the tendency of the near-wall anisotropy towards the two-component limit. © 2012 by M. Konopka, M. Meinke, and W. Schröder.

Hartmann A.,RWTH Aachen | Hartmann A.,Institute of Aerodynamics | Klaas M.,RWTH Aachen | Klaas M.,Institute of Aerodynamics | And 2 more authors.
AIAA Journal | Year: 2013

The influence of coupled heave and pitch oscillations on the transonic flowfield over a supercritical DRA 2303 laminar-type airfoil in buffet conditions is analyzed using a combination of time-resolved stereo particle-image velocimetry and unsteady pressure measurements. Three flow configurations at a Mach number of Ma∞= 0.73 and at an angle of attack of α = 3.5 deg are analyzed to show the impact of the fluid-structure interaction on the overall flow structure: the flow over a fixed DRA 2303, the flow over a DRA 2303 undergoing flow-induced motion, and the flow over a DRA 2303 experiencing a forced sinusoidal heave/pitch motion at frequencies in the range of the natural buffet frequency of the airfoil. The results are analyzed regarding the origin and nature of the unsteady shock-boundary-layer interaction, the fluid-structure response behavior, and the influence of the structural motion on the shock oscillation. For the buffet flow, the shock motion is determined by the sound pressure level generated at the trailing edge. At buffet conditions, the frequency of the self-sustained airfoil motion is in the same range as the natural buffet frequency. At forced oscillation at excitation frequencies in the buffet range, the shock oscillation locks into the excitation frequency indeed, but the sound-wave-based feedback loop is hardly influenced by the airfoil motion. Copyright © 2013 by Christopher Porter, R. Mark Rennie, Eric J. Jumper.

Nottebrock B.,Institute of Aerodynamics | Geurts K.J.,Institute of Aerodynamics | Schroder W.,Institute of Aerodynamics
AIAA AVIATION 2014 -7th AIAA Flow Control Conference | Year: 2014

An adverse pressure gradient turbulent boundary layer has been established to show that the micro-pillar shear-stress sensor can be used in non-constant free-stream flows to measure the wall-shear stress distribution. The quality of the reference has been ensured by additional measurements such as pressure and PIV measurements, and large-eddy simulations. The momentum based Reynolds number is Reθ = 4420 and the Clauser parameter characterizing the pressure gradient is β = 1.89. An introductory discussion on the sensitivity of the micro-pillar shear-stress sensor reveals only little effect of a pressure gradient on the wall-shear stress results since the dependence of the drag force on the velocity distribution is damped. Hence, the sensor is only little sensitive to discrepancies from the linear law of the wall due to a positive pressure gradient. Measurements of wall-shear stress in an APG TBL using MPS3 are presented. The behavior and accuracy of the sensor is analyzed using the statistical moments. They coincide with the results of the numerical simulation. Furthermore, they reveal an effect of the pressure gradient on the wall-shear stress. Turbulence intensity, skewness and kurtosis are smaller compared to the zero-pressure gradient turbulent boundary layer. This confirms the decrease in number and strength of the near-wall high-speed streaks in the APG TBL which is in good agreement with the findings of Harun et al.

Eitel-Amor G.,Institute of Aerodynamics | Meinke M.,Institute of Aerodynamics | Schroder W.,Institute of Aerodynamics
20th AIAA Computational Fluid Dynamics Conference 2011 | Year: 2011

The present study focuses on the implementation of a local grid-refinement technique into a single-relaxation-time Lattice-Boltzmann Method (LBM). To demonstrate the performance of this method, the flow past a circular cylinder at Reynolds numbers ReD = 20, ReD = 40, and ReD = 100 and the flow past a sphere at ReD = 100, ReD = 300, ReD = 3700, and ReD = 104 are simulated. At turbulent flows the large-eddy simulation concept is used. The LBM plus local grid refinement yields accurate temporal and spatial results and increases the computational efficiency due to a drastic reduction of the cell number. In two dimensions the computational effort is reduced by a factor of more than 50 and in three dimensions a reduction by a factor above 400 is achieved. © 2011 by Institute of Aerodynamics.

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